Turbine vane cooling arrangement

ABSTRACT

A vane includes a pair of airfoils having a plurality of film cooling holes that extend through an exterior surface of the airfoils, the plurality of film cooling holes including at least a first subset of film cooling holes, wherein the first subset of film cooling holes break through the exterior surface at geometric coordinates set forth herein. Each of the geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto a turbine vane that may be incorporated into a gas turbine engine.The vane's airfoils and platforms include a plurality of film coolingholes as part of a cooling arrangement.

Gas turbine engines typically include a compressor section, a combustorsection and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

Both the compressor and turbine sections may include alternating seriesof rotating blades and stationary vanes that extend into the core flowpath of the gas turbine engine. For example, in the turbine section,turbine blades rotate and extract energy from the hot combustion gasesthat are communicated along the core flow path of the gas turbineengine. The turbine vanes, which generally do not rotate, guide theairflow and prepare it for the next set of blades.

Each of the blades and vanes include airfoils that extend into the coreflow path of the gas turbine engine between inner and outer platforms.Cooling airflow is communicated into an internal core of the airfoil andcan be discharged through the plurality of film cooling holes to providea boundary layer of film cooling air along the external surface of theairfoil and platforms. The film cooling air provides a barrier thatprotects the underlying substrate of the vane from the hot combustiongases that are communicated within the core flow path.

SUMMARY

In one exemplary embodiment a vane includes a pair of airfoils having aplurality of film cooling holes that extend through an exterior surfaceof the airfoils, the plurality of film cooling holes including at leasta first subset of film cooling holes, wherein the first subset of filmcooling holes break through the exterior surface at geometriccoordinates in accordance with Cartesian coordinate values of X, Y and Zas set forth in Table 1, wherein each of the geometric coordinates ismeasured from a reference point on a leading edge rail of a platform ofthe vane.

In another example of the above described vane the Cartesian coordinatevalues of Table 1 are expressed in inches.

In another example of any of the above described vanes the referencepoint includes a pin hole of the platform.

In another example of any of the above described vanes the plurality offilm cooling holes are spaced along a span of the airfoil body inmultiple collinearly aligned rows.

In another example of any of the above described vanes the first subsetof film cooling holes are disposed on a pressure side filet of theairfoil body.

Another example of any of the above described vanes further includes atleast a second subset of film cooling holes disposed on a pressure side,a suction side and a leading edge of the airfoil body.

In another example of any of the above described vanes the airfoilsextend between inner and outer platforms, the inner and outer platformsinclude a second plurality of film cooling holes that extend through anexterior surface of the platforms.

In another example of any of the above described vanes the plurality offilm cooling holes included in the first subset of film cooling holesare the only cooling holes disposed within a radially outward filetedregion of a pressure side of each airfoil.

In another example of any of the above described vanes the vane is afirst turbine stage vane.

An exemplary method for cooling a vane doublet includes generating afilm cooling layer across a pressure side filet of each airfoil in thevane doublet by passing air through a plurality of film cooling holesthat extend through an exterior surface of the airfoils, the pluralityof film cooling holes including at least a first subset of film coolingholes, wherein the first subset of film cooling holes break through theexterior surface at geometric coordinates in accordance with Cartesiancoordinate values of X, Y and Z as set forth in Table 1, wherein each ofthe geometric coordinates is measured from a reference point on aleading edge rail of a platform of the vane.

Another example of the above described method for cooling a vane doubletfurther includes generating a film cooling layer across at least one ofa pressure side, a suction side and a leading edge of the airfoil bodyby passing air through a plurality of film cooling holes disposed acrossthe at least one of the pressure side, the suction side and the leadingedge.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 illustrates a schematic, cross-sectional view of a gas turbineengine.

FIG. 2 schematically illustrates a doublet stator vane that can beincorporated into a gas turbine engine.

FIG. 3 schematically illustrates another view of the vane of FIG. 2.

FIG. 4 schematically illustrates a cross-sectional view of one of anairfoil that includes a plurality of film cooling holes as part of anairfoil cooling arrangement of the vane.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmenter section (not shown) among other systems for features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited toturbofan engines and these teachings could extend to other types ofengines, including but not limited to, three-spool engine architectures.

The gas turbine engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centerlinelongitudinal axis A. The low speed spool 30 and the high speed spool 32may be mounted relative to an engine static structure 33 via severalbearing systems 31. It should be understood that additional bearingsystems 31 may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 34 thatinterconnects a fan 36, a low pressure compressor 38 and a low pressureturbine 39. The high speed spool 32 includes an outer shaft 35 thatinterconnects a high pressure compressor 37 and a high pressure turbine40. In this embodiment, the inner shaft 34 and the outer shaft 35 aresupported at various axial locations by bearing systems 31 positionedwithin the engine static structure 33.

A combustor 42 is arranged between the high pressure compressor 37 andthe high pressure turbine 40. A mid-turbine frame 44 may be arrangedgenerally between the high pressure turbine 40 and the low pressureturbine 39. The mid-turbine frame 44 supports one or more bearingsystems 31 of the turbine section 28. The mid-turbine frame 44 mayinclude one or more airfoils 46 that may be positioned within the coreflow path C.

The inner shaft 34 and the outer shaft 35 are concentric and rotate viathe bearing systems 31 about the engine centerline longitudinal axis A,which is co-linear with their longitudinal axes. The core airflow iscompressed by the low pressure compressor 38 and the high pressurecompressor 37, is mixed with fuel and burned in the combustor 42, and isthen expanded over the high pressure turbine 40 and the low pressureturbine 39. The high pressure turbine 40 and the low pressure turbine 39rotationally drive the respective high speed spool 32 and the low speedspool 30 in response to the expansion.

In some non-limiting examples, the gas turbine engine 20 is ahigh-bypass geared aircraft engine. In a further example, the gasturbine engine 20 bypass ratio is greater than about six (6:1). Theexample gas turbine engine 20 can be a geared turbofan engine thatincludes an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3, and in another example is greater thanabout 2.5:1. The geared turbofan enables operation of the low speedspool 30 at higher speeds which can increase the operational efficiencyof the low pressure compressor 38 and low pressure turbine 39 and renderincreased pressure in a fewer number of stages.

The low pressure turbine 39 pressure ratio is pressure measured prior tothe inlet of the low pressure turbine 39 as related to the pressure atthe outlet of the low pressure turbine 39 prior to an exhaust nozzle ofthe gas turbine engine 20. In one non-limiting embodiment, the bypassratio of the gas turbine engine 20 is greater than about ten (10:1), thefan diameter is significantly larger than that of the low pressurecompressor 38, and the low pressure turbine 39 has a pressure ratio thatis greater than about 5 (5:1). It should be understood, however, thatthe above parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.55. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45. In anothernon-limiting embodiment the low fan pressure ratio is from 1.1 to 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second (350.5 meters/second).

Each of the compressor section 24 and the turbine section 28 may includealternating rows of rotor assemblies and vane assemblies (shownschematically) that carry one or more airfoils that may extend into thecore flow path C. For example, the rotor assemblies can carry aplurality of rotating blades 25, while each vane assembly can carry aplurality of vanes 27 that extend into the core flow path C. The blades25 of the rotor assemblies create or extract energy (in the form ofpressure) from core airflow that is communicated through the gas turbineengine 20. The vanes 27 of the vane assemblies direct core airflow tothe blades 25 of the rotor assemblies to either add or extract energy.

Various components of the gas turbine engine 20, including airfoils ofthe compressor section 24 and the turbine section 28, may be subjectedto repetitive thermal cycling under widely ranging temperatures andpressures. The hardware of the turbine section 28 is particularlysubjected to relatively extreme operating conditions. Therefore, somecomponents may require airfoil cooling arrangements for cooling theairfoils that extend into the core flow path C. Exemplary airfoilcooling arrangements that include internal cooling circuits and filmcooling holes are described herein.

FIGS. 2 and 3 illustrate a doublet stator vane 50 that may beincorporated into a gas turbine engine, such as the gas turbine engine20. A “doublet” includes a pair of airfoils joined to the same inner andouter platforms. The vane 50 of this particular embodiment is a firststage turbine vane of the turbine section 28. However, this disclosureis not limited to this particular vane and could extend to any vane thatis disposed within the core flow path C of the gas turbine engine 20.

The vane 50 includes two airfoils 52 that extend between an innerplatform 54 (on an inner diameter side) and an outer platform 56 (on anouter diameter side). Each airfoil 52 includes a leading edge 58, atrailing edge 60, a pressure side 62 and a suction side 64. Each airfoil52, including the pressure side 62 and the suction side 64, extends inchord Cx between the leading edge 58 and the trailing edge 60 andextends in span S between the inner platform 54 and the outer platform56.

A gas path 70 is communicated axially downstream through the gas turbineengine 20 in a direction that extends from the leading edge 58 towardthe trailing edge 60 of the airfoil 52. The gas path 70 (for thecommunication of core airfoil along the core flow path C) extendsbetween an inner gas path 72 associated with the inner platform 54 andan outer gas path 74 associated with the outer platform 56 of the vane50. The inner platform 54 and the outer platform 56 are connected to theairfoils 52 at the inner and outer gas paths 72, 74 via fillets 75.

Both the inner platform 54 and the outer platform 56 include leadingedge rails 66 and trailing edge rails 68 having one or more engagementfeatures 69 for mounting the vane 50 to the gas turbine engine 20, suchas to an engine casing. Other engagement feature configurations arecontemplated as within the scope of this disclosure, including but notlimited to, hooks, rails, bolts, rivets, tabs and/or other features thatcan be incorporated into the vane 50 to retain the vane 50 to the gasturbine engine 20. In this exemplary embodiment, the leading edge rail66 of the inner platform 54 includes a pin hole 82 having a center point84 (See FIG. 2).

Referring to FIG. 4, the airfoils 52 include an airfoil coolingarrangement that can include an internal cooling circuit 76 and aplurality of film cooling holes 78 that extend through an exteriorsurface 53 of the airfoil 52. In the illustrated example, the filets 75on the pressure side of each vane 50 are exposed to heat relatedstresses and a subset of the film cooling holes 78 are positioned on thepressure side filets 75. The internal cooling circuit 76 can receive acooling airflow CA to cool the internal surfaces of the airfoil 50 (SeeFIG. 3). In one exemplary embodiment, the cooling airflow CA is a bleedairflow that can be sourced from the compressor section 24 or any otherportion of the gas turbine engine 20 that is upstream from the vane 50.The internal cooling circuit 76 may include one or more cavities 80 thatdefine hollow openings within each airfoil 52. The cooling airflow CAcan be communicated through the cavities 80, which extend across anentire length of the airfoil 52, to cool the internal surfaces of theairfoil 52.

The plurality of film cooling holes 78 of the airfoil coolingarrangement can be formed through the airfoil 52 (between the exteriorsurface 53 and one or more of the cavities 80) such that each filmcooling hole 78 breaks out through the exterior surface 53 of theairfoil 52. In this exemplary embodiment, the filets 75 on the pressureside 62 of each vane include film cooling holes 78. In addition, filmcooling holes can be included at other locations of the leading edge 58,the pressure side 62 and the suction side 64. The film cooling holes 78may embody a cone shape or a round shape. Other shapes are alsocontemplated as within the scope of this disclosure.

FIG. 4 illustrates a cross-sectional view of an exemplary airfoil 50.The internal cooling circuit 76 of the airfoil 50 includes multiplecavities 80A-80C that can receive cooling airflow CA to cool theinternal surfaces of the airfoil 50. In alternative configurations,alternative internal cooling cavity configurations can be utilized andachieve similar results. The plurality of film cooling holes 78 are influid communication with one or more of the cavities 80A-80C. Coolingairflow CA can be communicated into and through the cavities 80A-80C andcan then be discharged through the plurality of film cooling holes 78 toprovide a boundary layer of film cooling air along the exterior surface53 of the airfoil 50. The film cooling air may provide a barrier thatprotects the underlying substrate of the airfoil 50 from the hotcombustion gases that are communicated within the core flow path C.

The plurality of film cooling holes 78 are spaced apart along the span Sof the airfoil 52 for discharging the cooling airflow CA and providing aboundary layer of film cooling air along the exterior surface 53 of theairfoil 52. In this exemplary embodiment, with reference to FIG. 4 andTable 1, the pressure side 62 includes PWA, PWB, PWC, PWD, etc. of filmcooling holes 78 of film cooling holes 78. Additional holes can beprovided but are not shown for clarity. In this disclosure, holesidentified with the letter “P” refer to the rows of the pressure side62. The locations shown in FIG. 4 are schematic.

The breakout point of each film cooling holes 78 refers to the geometriclocation along the vane 50 at which the film cooling hole centerlinebreaks through or protrudes out of the exterior surface of the vane 50.The breakout points of each of the plurality of film cooling holes 78can be described in terms of sets of Cartesian coordinates, provided inTable 1 (airfoils), defined along x, y and z axes as measured from aspecific reference point of the vane 50, as is further discussed below.As shown in FIGS. 2 and 3 (with continued reference to FIG. 1), the xaxis is defined along the direction of the engine centerlinelongitudinal axis A, the y axis is defined in a substantiallycircumferential or rotational direction about the engine centerlinelongitudinal axis A, and the z axis is defined in a radial directionthat is substantially perpendicular to the engine centerlinelongitudinal axis A.

In the exemplary embodiment, each of the geometric coordinates aremeasured from a pin hole 82 of the leading edge rail 66 of the innerplatform 54 of the vane 50. By measuring the geometric coordinates (interms of x, y and z values) from a center point 84 of the pin hole 82,the external break through points of a subset of the film cooling holes78 of the airfoil 52. In the exemplary embodiment, the subset of filmcooling holes 78 are the only film cooling holes breaking out in theradially outward filet 75 of the pressure side of each airfoil.Additional film cooling holes 78 are disposed across the surface of theairfoil. The Table 1 values for the x, y and z coordinates represent thetrue position of a nominal part and are listed in inches in thisembodiment. However, the values of the Tables could be converted tomillimeters by multiplying by 25.4, or could be converted to any otherunits. In other words, due to manufacturing tolerances, the externalbreakout of the centerline of each film cooling hole 78 can fall withina 0.200 inch diameter circle enscribed on the surface of the airfoil 50.

Table 1 identify each film cooling hole 78 by assigning a unique,three-letter identifier to each film cooling hole 78. The first twoletters of the three-letter identifier identify the row (PW) where thefilm cooling hole is located. The third letter corresponds to thespecific hole number of a particular row identified by the first twoletters. The film cooling holes of each group are numbered from back tofront, right airfoil to left airfoil, with the letter A representing thefilm cooling hole 78 closest to the trailing edge 60 on the rightairfoil 52, B representing the film cooling hole 78 closest to theleading edge 58 on the right airfoil 52, and subsequent letters assignedto the film cooling holes 78 in the same manner on the left airfoil(i.e., C, D).

TABLE 1 airfoil cooling holes Hole Name X Y Z PWA 0.617-1.0171.542-1.942 2.200-2.600 PWB 0.445-0.845 1.329-1.729 2.292-2.692 PWC0.617-1.017 −0.290-0.110   2.263-2.663 PWD 0.445-0.845 −0.517-−0.1172.313-2.713

The hole locations defined in Table 1 are defined at ambient,non-operating, conditions on a coated doublet stator vane 50.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

In further examples, additional film cooling holes 78 can be includeddisposed across gaspath surfaces of the inner and outer platforms 54, 56to provide similar film cooling benefits to the platform bodies.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that various modifications could come within the scope of thisdisclosure. For these reasons, the following claims should be studied todetermine the true scope and content of this disclosure.

What is claimed is:
 1. A vane comprising: a pair of airfoils having a plurality of film cooling holes that extend through an exterior surface of said airfoils, the plurality of film cooling holes including at least a first subset of film cooling holes, wherein said first subset of film cooling holes break through said exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1, wherein each of said geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.
 2. The vane as recited in claim 1, wherein said Cartesian coordinate values of Table 1 are expressed in inches.
 3. The vane as recited in claim 1, wherein said reference point includes a pin hole of said platform.
 4. The vane as recited in claim 1, wherein said plurality of film cooling holes are spaced along a span of said airfoil body in multiple collinearly aligned rows.
 5. The vane as recited in claim 1, wherein said first subset of film cooling holes are disposed on a pressure side filet of said airfoil body.
 6. The vane as recited in claim 5, further comprising at least a second subset of film cooling holes disposed on a pressure side, a suction side and a leading edge of said airfoil body.
 7. The vane as recited in claim 1, wherein the airfoils extend between inner and outer platforms, the inner and outer platforms include a second plurality of film cooling holes that extend through an exterior surface of said platforms.
 8. The vane as recited in claim 1, wherein the plurality of film cooling holes included in the first subset of film cooling holes are the only cooling holes disposed within a radially outward fileted region of a pressure side of each airfoil.
 9. The vane as recited in claim 1, wherein the vane is a first turbine stage vane.
 10. A method for cooling a vane doublet, comprising: generating a film cooling layer across a pressure side filet of each airfoil in the vane doublet by passing air through a plurality of film cooling holes that extend through an exterior surface of said airfoils, the plurality of film cooling holes including at least a first subset of film cooling holes, wherein said first subset of film cooling holes break through said exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1, wherein each of said geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.
 11. The method of claim 10, further comprising generating a film cooling layer across at least one of a pressure side, a suction side and a leading edge of said airfoil body by passing air through a plurality of film cooling holes disposed across the at least one of the pressure side, the suction side and the leading edge. 